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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also completed fabrication of several subscale and full-scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

DED↗
NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also completed fabrication of several subscale and full-scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

Additive Manufacturing↗
NASA NTRSDate not supplied

Why Developing A Full-Sun UV Spectrograph Sounding Rocket Is Difficult but Necessary

For a conference focused on a 20-year perspective of where we have been and where we are going, this talk will highlight specifically why we need a sounding rocket to lead the way in obtaining high-resolution UV spectroscopy of our sun-as-a-star. The currently available data is either sparse or at a much lower resolution when compared with Hubble (HST) spectra of other stars. Because of this, we cannot say with certainty if our Sun’s variability is typical. This limitation is hindering our quest to place our Sun in it’s rightful place on our star-charts. Ultraviolet light is a key component in determining this. Elements such as Hydrogen and Iron have strong emission lines here, and from them we can determine many things about the star from which they originate. The width of these spectral features can tell us about the relative plasma activity on the surface of the star, the abundance of those elements, and of course the temperatures. This has of course been done with many instruments that have very narrow field of views and/or narrow bandwidths. The only way to directly compare our Sun’s UV ”fingerprint” with other stars is to take an image of it in the same manner by which HST does: as a point source object. Our team’s aim has been to develop a sounding rocket capable of doing just that. Along the way, the many challenges we face are producing many solutions. A unique take on the traditional Rowland circle spectrograph will allow us to cover a large range of UV spectra with minimal moving parts. A novel in-situ gain calibration technique will be used in order to maintain the high level of radiometric accuracy desired. Naturally, complications arise when one takes UV measurements as you are flying out of earth’s atmosphere and back into it. Though it may be minor, the absorbed UV spectra from the atmosphere will have to be accounted for. In preparing for this, we have stumbled across a way in which those ”telluric” absorption lines may prove useful for spectral calibration purposes. In addition, raw data from previous sounding rocket missions may have hidden within them useful information for the atmospheric-physics community. In this talk, we will highlight the unique calibration problems and solutions we have faced, and how those are of greater importance to the greater heliophysics and atmospheric communities. With these amazing developments at hand, the future of heliophysics instrumentation is looking bright. This material is based upon work supported by the NSF EPSCoR RII-Track-1.2a (Non-invasive plasma diagnostics for LTP) Cooperative Agreement OIA-1655280. Any opinions, findings, and conclusions or recommendations expressed in this material are those of the authors and do not necessarily reflect the views of the National Science Foundation.

NASA, UAH, MSU, Sounding Rocket, Aeronomy, Instrum↗
DOE OSTI2024

Satellite-Based Assessment of Rocket Launch and Coastal Change Impacts on Cape Canaveral Barrier Island, Florida, USA

The Cape Canaveral Barrier Island, home to the National Aeronautics and Space Administration (NASA)’s Kennedy Space Center and the United States (U.S.) Space Force’s Cape Canaveral Space Force Station, is situated in a unique ecological transition zone that supports diverse wildlife. This study evaluates the recent changes in vegetation cover (2016–2023) and dune elevation (2007–2017) within the Cape Canaveral Barrier Island using high-resolution optical satellite and light detection and ranging (LiDAR) data. The study period was chosen to depict the time period of a recent increase in rocket launches. The study objectives include assessing changes in vegetation communities, identifying detectable impacts of liquid propellant launches on nearby vegetation, and evaluating dune elevation and tide level shifts near launchpads. The results indicate vegetation cover changes, including mangrove expansion in wetland areas and the conversion of coastal strands to denser scrubs and hardwood forests, which were likely influenced by mild winters and fire management. While detectable impacts of rocket launches on nearby vegetation were observed, they were less severe than those caused by solid rocket motors. Compounding challenges, such as rising tide levels, beach erosion, and wetland loss, potentially threaten the resilience of launch operations and the surrounding habitats. The volume and scale of launches continue to increase, and a balance between space exploration and ecological conservation is required in this biodiverse region. This study focuses on the assessment of barrier islands’ shorelines.

54 ENVIRONMENTAL SCIENCES↗
DOE OSTI2024

Simulations of the collection of mesospheric dust particles with a rocket instrument

We investigate the collection of dust particles in the mesosphere with the MESS (MEteoric Smoke Sampler) instrument that is designed to fly on a sounding rocket. We assume that the ice particles that form in the polar mesosphere between 80 and 85 km altitude in summer contain meteoric smoke particles; and these should be collected with MESS. The instrument consists of a collection device with an opening and closure mechanism, as well as an attached conic funnel which increases the sampling area in comparison to the collection area. Dust particles are collected either directly after passing through the instrument or indirectly after colliding with and fragmenting on the funnel wall. We calculate the dust and fragment trajectories in the detector to determine the collection efficiency for different particle sizes, rocket velocities, and heights, and we find the final velocities and the temperatures of the particles. The considered design has a sampling area of 62.78 mm diameter and a collection area of 20 mm diameter. For the conditions at the rocket launch site in Andøya, Norway, we estimate the collection of meteoric smoke particles contained in the ice particles to be ~10 12 –10 14 amu mm –2 . The estimated temperatures suggest that the composition of these smoke particles is not affected by the collection. Our calculations also show that keeping the instrument open above 85 km altitude increases the amount of small smoke particles that are directly collected. The directly collected smoke particles are heated as they decelerate, which can affect their composition.

47 OTHER INSTRUMENTATION↗
NASA NTRS1949

Experimental Study of Gas-Flow Separation in Overexpanded Exhaust Nozzles for Rocket Motors

In a rocket motor having a de Laval nozzle with a fixed expansion ratio, there is only one pressure ratio (p(sub c)/p(sub o)) for which this expansion ratio is correct. At any other pressure ratio the gases will be either underexpanded or overexpanded, depending on whether the pressure ratio is higher or lower than the correct value. Rocket motor flame patterns illustrating the four possible regimes are shown in Figures 1 to 4, inclusive. In Figure 1, the pressure ratio is higher than the correct one for the fixed expansion ratio of the nozzle, resulting in underexpansion of the gases. in Figure 2, the pressure ratio is the correct one. In Figure 3, the pressure ratio is only slightly lower than the correct one, resulting in overexpansion without separation of the gases from the wall. In Figure 4, the pressure ratio is considerably below the correct one, resulting in overexpansion with separation of the gases from the nozzle wall. The present report is concerned primarily with this fourth type of flow, that is, overexpansion of the gases with separation of the flow from the wall. Such a situation exists in a rocket motor operating at a considerably lower altitude than that for which its exhaust nozzle was designed.

Cowles, Frederick B.↗
NASA NTRS1968

Description and Flight Performance Results of the WASP Sounding Rocket

A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

De Pauw, J. F.↗
NASA NTRS1971

The United States sounding rocket program

The United States sounding rocket program is discussed. The program is concerned with the fields of solar physics, galactic astronomy, fields and particles, ionospheric physics, aeronomy, and meteorology. Sounding rockets are described with respect to propulsion systems, gross weight, and capabilities. Instruments used to conduct ionospheric probing missions are examined. Results of previously conducted sounding rocket missions are included.

Source record↗
NASA NTRS1972

Pulse performance analysis for small hypergolic-propellant rocket engines

Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 1 to 100 lbs were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical and test performance. The theoretical analysis is modified to obtain a semi-empirical model for hypergolic propellant rockets which is demonstrated to be reasonably accurate for two different engine configurations over a considerable range of duty cycles.

Smith, G. W.↗
NASA NTRS1972

The geometry and physical properties of exhaust clouds generated during the static firing of S-1C and S-2 rocket engines

An experimental program was conducted during the static firing of the S-1C stage 13, 14, and 15 rocket engines and the S-2 stage 13, 14, and 15 rocket engines. The data compiled during the experimental program consisted of photographic recordings of the time-dependent growth and diffusion of the exhaust clouds, the collection of meteorological data in the ambient atmosphere, and the acquisition of data on the physical structure of the exhaust clouds which were obtained by flying instrumented aircraft through the clouds. A new technique was developed to verify the previous measurements of evaporation and entrainment of blast deflector cooling water into the cloud. The results of the experimental program indicate that at the lower altitudes the rocket exhaust cloud or plume closely resembles a free-jet type of flow. At the upper altitudes, where the cloud is approaching an equilibrium condition, structure is very similar to a natural cumulus cloud.

Forbes, R. E.↗
NASA NTRS1972

Liquid rocket performance computer model with distributed energy release

Development of a computer program for analyzing the effects of bipropellant spray combustion processes on liquid rocket performance is described and discussed. The distributed energy release (DER) computer program was designed to become part of the JANNAF liquid rocket performance evaluation methodology and to account for performance losses associated with the propellant combustion processes, e.g., incomplete spray gasification, imperfect mixing between sprays and their reacting vapors, residual mixture ratio striations in the flow, and two-phase flow effects. The DER computer program begins by initializing the combustion field at the injection end of a conventional liquid rocket engine, based on injector and chamber design detail, and on propellant and combustion gas properties. It analyzes bipropellant combustion, proceeding stepwise down the chamber from those initial conditions through the nozzle throat.

Combs, L. P.↗
NASA NTRS1972

Results of a rocket-Nimbus sounder comparison experiment

Stratospheric temperature data obtained from instrumentation on board Nimbus 3 and 4 are compared with conventional rocket soundings to determine the compatibility of measurements. The Arcasonde 1A and Datasonde were the primary rocketsondes used, but acoustic grenade and pitot probe soundings were also involved. Observed temperature profiles and computed radiances obtained from rocket soundings are compared with 15-microns data provided by the satellite infrared spectrometers, the infrared interferometer spectrometer, and the selective chopper radiometer onboard the satellites. A reasonable agreement between the rocket and satellite data is indicated, but variations are noted which are related to the sounding techniques.

Miller, A. J.↗
NASA NTRS1971

Nozzle and cavity wall cooling limitations on specific impulse of a gas-core nuclear rocket.

Experimental and theoretical study of the performance of a gas-core nuclear rocket, showing that in a uranium plasma nuclear rocket with an 8-ft cavity diameter the cavity wall can be cooled up to a power level of 7400 MW when the rocket operates at a pressure of 1000 atm and the propellant mass flow rate is 10 lbm/sec. A maximum cavity specific impulse of 5800 sec was obtained under such operating conditions. The fact that the wall heat flux was much lower for reactor powers below this level is linked to the presence of a relatively cool, opaque insulating layer of seeded propellant between the hot plasma and the solid wall. The additional coolant required for nozzle protection reduced the maximum cavity specific impulse to 5200 sec.

Kascak, A. F.↗
NASA NTRS1971

Technology for low cost solid rocket boosters.

A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

Ciepluch, C.↗
NASA NTRS1972

Development of boron epoxy rocket motor chambers.

A 71 cm diameter 74 cm length boron/epoxy composite rocket motor chamber was designed based on the geometric configuration of the JPL Applications Technology Satellite titanium alloy apogee motor chamber. Because analyses showed large stress concentrations in the domes, the configuration was modified using the same basic constraints for openings and attachments. The rocket motor chamber was then fabricated by filament winding with boron/epoxy tape and hydrostatically tested to failure at 264 N/sq cm, 57.2 N/sq cm above the design value. Two more rocket motor chambers were fabricated with the same basic constraints, but shortened to 57.6 cm for a smaller propellant load. The first of these short chambers failed in proof because of filament winding fabrication difficulties. The second chamber was successfully fabricated and passed the hydrostatic proof test.

Jensen, W. M.↗
NASA NTRS1972

Pulse performance analysis for small hypergolic propellant rocket engines

Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 5 to 445 N (1 to 100 lb) were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical performance and test performance. The theoretical analysis is modified to obtain a semiempirical model for hypergolic propellant rockets.

Smith, G. W.↗
NASA NTRS1972

A program of high resolution X-ray astronomy using sounding rockets

Two Aerobee 170 sounding rocket payloads were flown at the White Sands Missile Range: (1) a focusing X-ray collector on 31 March 1972; and (2) a high resolution telescope on 4 August 1972. Data has been reduced from each of these flights. In the first flight both the rocket and the experiment instrumentation performed adequately, and it is clear that at least the minimum scientific objectives were attained. In the second flight the attitude control system failed to point the telescope at the target for a sufficient length of time. However examination of final preflight checkout data and some flight data indicate that the instrumentation for this rocket payload was functioning according to expectations.

Source record↗
NASA NTRS1972

Study of solid rocket motor for space shuttle booster, volume 2, book 2

A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.

Source record↗